Health monitoring for hollow blades

ABSTRACT

A system for detecting a defect in a hollow blade includes a pressure sensor mounted within a cavity of the hollow blade, wherein the pressure sensor measures a pressure within the cavity and provides an output indicative of measured pressure; and a controller to receive the output from the pressure sensor, wherein the controller indicates a defect in the hollow blade based upon the received output.

BACKGROUND

The present invention is related to gas turbine engines, and in particular to health monitoring of hollow blades.

Gas turbine engines include a fan section, a compressor section, a combustion section, and a turbine section. The fan section includes a rotor assembly and a stator assembly. The rotor assembly includes a plurality of fan blades. Fan blades may, in order to reduce weight and provide other advantages, be implemented as hollow fan blades. These fan blades, including hollow fan blades, are susceptible to defects during operation. Mechanical defects in components, such as cracks in the fan blade, must be detected in order to avoid failure of the component. In the past, visual inspection has been performed to detect cracks in fan blades. It is desirable to have a system that provides real-time health monitoring of hollow fan blades.

SUMMARY

A system for detecting a defect in a hollow blade includes a pressure sensor and a controller. The pressure sensor is mounted within a cavity of the hollow blade, measures a pressure within the cavity, and provides an output indicative of measured pressure. The controller receives the output from the pressure sensor and indicates a defect in the hollow blade based upon the received output.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically illustrates an example gas turbine engine that includes a fan section, a compressor section, a combustor section and a turbine section.

FIG. 2A schematically illustrates a hollow fan blade with an internal pressure sensor according to an embodiment of the present invention.

FIG. 2B schematically illustrates a cross section of a hollow fan blade with an internal pressure sensor according to an embodiment of the present invention.

FIG. 3 is a flowchart illustrating a method of detecting a crack in a hollow fan blade according to an embodiment of the present invention.

DETAILED DESCRIPTION

The present invention describes a system and method for monitoring the health of a hollow blade. The system includes a pressurized gas within the hollow blade, a pressure sensor within the hollow blade, and a controller. The hollow blade is filled with a pressurized gas, such as air, CO₂, Nitrogen, or any other inert gas. The hollow blade may also be filled with vacuum. The pressure sensor is mounted on an internal boundary of the blade, or at least a portion of the pressure sensor is in fluid communication with the hollow cavity. The pressure sensor measures the pressure within the hollow cavity and detects a change in pressure. When the pressure sensor determines a change in pressure has occurred, the pressure sensor sends an indication to a controller, such as a full authority digital engine controller (FADEC). The FADEC sends an indication of the defect in the blade to, for example, the pilot of an aircraft. The pilot may then take immediate action, such as shutting down the engine that includes the defective blade.

FIG. 1 schematically illustrates an example gas turbine engine 20 that includes fan section 22, compressor section 24, combustor section 26 and turbine section 28. Alternative engines might include an augmenter section (not shown) among other systems or features. Fan section 22 drives air along bypass flow path B while compressor section 24 draws air in along core flow path C where air is compressed and communicated to combustor section 26. In combustor section 26, air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through turbine section 28 where energy is extracted and utilized to drive fan section 22 and compressor section 24.

Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section.

The example engine 20 generally includes low speed spool 30 and high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.

Low speed spool 30 generally includes inner shaft 40 that connects fan 42 and low pressure (or first) compressor section 44 to low pressure (or first) turbine section 46. Inner shaft 40 drives fan 42 through a speed change device, such as geared architecture 48, to drive fan 42 at a lower speed than low speed spool 30. High-speed spool 32 includes outer shaft 50 that interconnects high pressure (or second) compressor section 52 and high pressure (or second) turbine section 54. Inner shaft 40 and outer shaft 50 are concentric and rotate via bearing systems 38 about engine central longitudinal axis A.

Combustor 56 is arranged between high pressure compressor 52 and high pressure turbine 54. In one example, high pressure turbine 54 includes at least two stages to provide a double stage high pressure turbine 54. In another example, high pressure turbine 54 includes only a single stage. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine.

The example low pressure turbine 46 has a pressure ratio that is greater than about 5. The pressure ratio of the example low pressure turbine 46 is measured prior to an inlet of low pressure turbine 46 as related to the pressure measured at the outlet of low pressure turbine 46 prior to an exhaust nozzle.

Mid-turbine frame 58 of engine static structure 36 is arranged generally between high pressure turbine 54 and low pressure turbine 46. Mid-turbine frame 58 further supports bearing systems 38 in turbine section 28 as well as setting airflow entering low pressure turbine 46.

The core airflow C is compressed by low pressure compressor 44 then by high pressure compressor 52 mixed with fuel and ignited in combustor 56 to produce high speed exhaust gases that are then expanded through high pressure turbine 54 and low pressure turbine 46. Mid-turbine frame 58 includes vanes 60, which are in the core airflow path and function as an inlet guide vane for low pressure turbine 46. Utilizing vane 60 of mid-turbine frame 58 as the inlet guide vane for low pressure turbine 46 decreases the length of low pressure turbine 46 without increasing the axial length of mid-turbine frame 58. Reducing or eliminating the number of vanes in low pressure turbine 46 shortens the axial length of turbine section 28. Thus, the compactness of gas turbine engine 20 is increased and a higher power density may be achieved.

The disclosed gas turbine engine 20 in one example is a high-bypass geared aircraft engine. In a further example, gas turbine engine 20 includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). The example geared architecture 48 is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3.

In one disclosed embodiment, gas turbine engine 20 includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of low pressure compressor 44. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines.

A significant amount of thrust is provided by bypass flow B due to the high bypass ratio. Fan section 22 of engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft., with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point.

“Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.50. In another non-limiting embodiment the low fan pressure ratio is less than about 1.45.

“Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/518.7]^(0.5). The “Low corrected fan tip speed”, as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second.

The example gas turbine engine includes fan 42 that comprises in one non-limiting embodiment less than about twenty-six fan blades. In another non-limiting embodiment, fan section 22 includes less than about twenty fan blades. Moreover, in one disclosed embodiment low pressure turbine 46 includes no more than about six turbine rotors schematically indicated at 34. In another non-limiting example embodiment low pressure turbine 46 includes about three turbine rotors. A ratio between number of fan blades 42 and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine 46 provides the driving power to rotate fan section 22 and therefore the relationship between the number of turbine rotors 34 in low pressure turbine 46 and number of blades 42 in fan section 22 disclose an example gas turbine engine 20 with increased power transfer efficiency.

FIG. 2A is a schematic diagram of system 110 for monitoring the health of hollow fan blade 114 by, for example, detecting crack 112 in hollow fan blade 114. Although the present embodiment has been described with reference to a fan blade, the invention can be used for other blades or vanes, such as hollow compressor blades or vanes, or utilized for rotor blades for helicopters. Hollow fan blade 114 includes pressure side 116, suction side 118, internal cavity 120, and root 122. Sensor 124 is mounted to an inner wall of internal cavity 120 on either the inner wall of pressure side 116, the inner wall of suction side 118, or close to root 122 in internal cavity 120. Sensor 124 is in fluid communication with cavity 120. While not visible in FIG. 2A, the locations of internal cavity 120 and sensor 124 are indicated by a dotted line. Pressure side 116 is not visible in FIG. 2A and is shown in FIG. 2B. Sensor 124 is mounted using any well known mounting method such as, for example, machining a flat portion of the inner wall and attaching sensor 124 to the flat using a known adhesive, or using any known mechanical attachment means. In one embodiment, sensor 124 includes transceiver 126 for communicating with controller 128. Controller 128 includes transceiver 130 for receiving a signal from sensor 124. Controller 128 is any electronic controller, such as a FADEC. Transceivers 126 and 130 are any transceivers capable of sending and receiving wireless signals, such as radio-frequency (RF) transceivers. In another embodiment, wired signals are utilized to communicate between sensor 124 and controller 128.

FIG. 2B is a schematic diagram illustrating a cross section of hollow fan blade 114 taken along line 2B-2B of FIG. 2A. While illustrated showing sensor 124 mounted to the inner wall of suction side 118, sensor 124 may be mounted anywhere within cavity 120, including pressure side 116. Cavity 120 may include ribs 132. While illustrated as including three longitudinal ribs 132, internal cavity 120 may include any number of ribs necessary to provide structural support for hollow fan blade 114.

Internal cavity 120 of hollow fan blade 114 contains a known pressure. The pressure of the gas included in cavity 120 may be different than the expected atmospheric pressure external to hollow fan blade 114 during normal system operation. Because this expected atmospheric pressure changes with altitude, the gas included in cavity 120 may be selected to have a pressure different than any expected atmospheric pressure during normal system operation. This expected atmospheric pressure may include a range of, for example, any expected atmospheric pressure between sea level and an altitude of 40,000 feet. This may be accomplished by filling cavity 120 with a gas such as air, CO₂, Nitrogen, or any other inert gas having a pressure higher than any expected atmospheric pressure external to hollow fan blade 114 during normal system operation. Cavity 120 may also contain a known pressure that is less than any expected atmospheric pressure, such as approximately zero, by filling cavity 120 with vacuum.

Hollow fan blade 114 is produced, for example, by joining two metal sheets using a process such as diffusion bonding. To fill cavity 120 with a gas or with vacuum, the two metal sheets are joined in an environment containing the desired gas or vacuum. Pressure sensor 124 is mounted to one of the two metal sheets prior to joining the two sheets to form hollow fan blade 114.

Pressure sensor 124 is used to measure the pressure within internal cavity 120 during normal system operation. Pressure sensor 124 is any known pressure sensor. If a defect such as a crack occurs in fan blade 114, external air will enter cavity 120, or gas in cavity 120 will leak out, and the pressure within cavity 120 will change. In one embodiment, internal cavity 120 may be filled with vacuum. Pressure sensor 124 may be, for example, an absolute pressure sensor that provides controller 128 an output indicative of pressure relative to vacuum. If a crack occurs in fan blade 114, ambient air will enter cavity 120 and provide pressure greater than approximately zero. If sensor 124 provides an output indicative of pressure that is not the expected pressure within cavity 120, such as approximately zero, controller 128 will detect a defect in fan blade 114, and provide an indication of the defect to, for example, a pilot of the aircraft that includes fan blade 114. The pilot may then take immediate action, such as shutting down the gas turbine engine that includes fan blade 114 prior to the defect becoming any worse. This allows system 110 to detect a crack in fan blade 114 immediately, prior to the crack becoming larger and causing further issues with fan blade 114.

In another embodiment, pressure sensor 124 detects a pressure different than the expected pressure within cavity 120. Upon detection of such a condition, pressure sensor 124 provides an output to controller 128 indicative of the detected difference in pressure. Controller 128 may then indicate a defect in hollow fan blade 114 to, for example, a pilot of the aircraft that includes fan blade 114.

FIG. 3 is a flowchart illustrating method 140 of detecting crack 112 in hollow fan blade 114. At step 142, an expected pressure in internal cavity 120 is determined by filling cavity 120 with a pressurized gas, such as air, CO₂, Nitrogen, or any other inert gas. This may be accomplished by joining a pressure side sheet with a suction side sheet within an environment that contains the desired air, CO₂, Nitrogen, or other inert gas. Alternatively, cavity 120 could be filled with vacuum, having a known approximate zero pressure. At step 144, sensor 124 measures the pressure within cavity 120. At step 146, it is determined if the pressure within cavity 120 is different than the predetermined pressure. If the pressure is different than the predetermined pressure, method 140 proceeds to step 148. If the pressure is approximately the same as the predetermined pressure, method 140 returns to step 144. At step 148, sensor 124 indicates a change in pressure to controller 126 and controller 126 indicates a defective fan blade to, for example, a pilot of the aircraft that includes fan blade 114.

The following are non-exclusive descriptions of possible embodiments of the present invention.

A system for detecting a defect in a hollow blade according to an exemplary embodiment of this disclosure, among other possible things includes: a pressure sensor mounted within a cavity of the hollow blade, wherein the pressure sensor measures a pressure within the cavity and provides an output indicative of measured pressure, and a controller to receive the output from the pressure sensor, wherein the controller indicates a defect in the hollow blade based upon the received output.

The system of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components.

The hollow blade is a hollow fan blade.

The cavity is pressurized to a level different from an expected atmospheric pressure external to the hollow blade.

The pressurized level is less than the expected atmospheric pressure.

The pressurized level is greater than the expected atmospheric pressure.

The pressure sensor communicates with the controller wirelessly.

The pressure sensor compares the measured pressure to an expected pressure, and wherein the output indicates whether the measured pressure differs from the expected pressure.

The controller compares the output to an expected pressure and detects the defect based upon a difference between the output and the expected pressure.

The controller indicates the defect in the hollow blade to a pilot of an aircraft that includes a gas turbine engine containing the hollow blade.

A method for monitoring a hollow blade according to an exemplary embodiment of this disclosure, among other possible things includes: determining an expected pressure within a cavity of the hollow blade, measuring a current pressure within the cavity of the hollow blade using a pressure sensor, wherein the pressure sensor provides an output indicative of the measured pressure, and indicating a defect in the hollow blade based upon an output of the pressure sensor, wherein the indication of the defect is made using a controller.

The method of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components.

The hollow blade is a hollow fan blade.

Indicating the detected defect to a pilot of an aircraft that includes a gas turbine engine containing the hollow fan blade.

Filling the cavity with a gas having a known pressure.

Indicating a defect if the controller determines the measured pressure is different than the expected pressure.

A hollow component of a gas turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes: a cavity filled with a gas having a known pressure, and a pressure sensor mounted within the cavity, wherein the pressure sensor measures a pressure within the cavity, and wherein the pressure sensor provides an output to an engine controller indicative of the measured pressure.

The hollow component of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components.

The known pressure is greater than an expected atmospheric pressure external to the hollow component.

The known pressure is less than an expected atmospheric pressure external to the hollow component.

The pressure sensor further determines if the measured pressure is different than the known pressure.

The hollow component is a hollow fan blade.

While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims. 

1. A system for detecting a defect in a hollow blade, the system comprising: a pressure sensor mounted within a cavity of the hollow blade, wherein the pressure sensor measures a pressure within the cavity and provides an output indicative of measured pressure; and a controller to receive the output from the pressure sensor, wherein the controller indicates a defect in the hollow blade based upon the received output.
 2. The system of claim 1, wherein the hollow blade is a hollow fan blade.
 3. The system of claim 1, wherein the cavity is pressurized to a level different from an expected atmospheric pressure external to the hollow blade.
 4. The system of claim 3, wherein the pressurized level is less than the expected atmospheric pressure.
 5. The system of claim 3, wherein the pressurized level is greater than the expected atmospheric pressure.
 6. The system of claim 1, wherein the pressure sensor communicates with the controller wirelessly.
 7. The system of claim 1, wherein the pressure sensor compares the measured pressure to an expected pressure, and wherein the output indicates whether the measured pressure differs from the expected pressure.
 8. The system of claim 1, wherein the controller compares the output to an expected pressure and detects the defect based upon a difference between the output and the expected pressure.
 9. The system of claim 1, wherein the controller indicates the defect in the hollow blade to a pilot of an aircraft that includes a gas turbine engine containing the hollow blade.
 10. A method for monitoring a hollow blade, the method comprising: determining an expected pressure within a cavity of the hollow blade; measuring a current pressure within the cavity of the hollow blade using a pressure sensor, wherein the pressure sensor provides an output indicative of the measured pressure; and indicating a defect in the hollow blade based upon an output of the pressure sensor, wherein the indication of the defect is made using a controller.
 11. The method of claim 10, wherein the hollow blade is a hollow fan blade.
 12. The method of claim 10, wherein indicating the defect in the hollow blade comprises indicating the detected defect to a pilot of an aircraft that includes a gas turbine engine containing the hollow blade.
 13. The method of claim 10, wherein determining an expected pressure within the cavity comprises filling the cavity with a gas having a known pressure.
 14. The method of claim 10, wherein indicating a defect based upon an output of the pressure sensor comprises indicating a defect if the controller determines the measured pressure is different than the expected pressure.
 15. A hollow component of a gas turbine engine, the hollow component comprising: a cavity filled with a gas having a known pressure; and a pressure sensor mounted within the cavity, wherein the pressure sensor measures a pressure within the cavity, and wherein the pressure sensor provides an output to an engine controller indicative of the measured pressure.
 16. The hollow component of claim 15, wherein the known pressure is greater than an expected atmospheric pressure external to the hollow component.
 17. The hollow component of claim 15, wherein the known pressure is less than an expected atmospheric pressure external to the hollow component.
 18. The hollow component of claim 15, wherein the pressure sensor further determines if the measured pressure is different than the known pressure.
 19. The hollow component of claim 15, wherein the hollow component is a hollow fan blade. 